Reaction motor employing intermittent explosive combustion and impulse turbine

ABSTRACT

A thrust producing gas turbine adapted for aircraft propulsion at both moderate and especially high thrust values, and characterized by intermittent explosive burning of fuel into the working fluid and the directing thereof into an impulse turbine for the partial absorption of energy by means of a decrease in velocity and for the conservation of pressure for continued and subsequent effective burning. The engine embodiment comprises a compressor section followed by a nozzle-burner section shaped for the explosive high pressure high velocity discharge of working fluid into the turbine section for the absorption of operating energy and conservation of pressure, and through which the working fluid exhausts through stator blades that reestablish axial flow and thrust; the working fluid completing its combustion and/or being injected with fuel for efficient afterburning. The compressor air is stratified as it passes through the nozzle-burner section for cooling effect and to supply the products of combustion for supporting complete primary and secondary burning.

United States Patent [1 1 Grow 1 1 REACTION MOTOR EMPLOYING INTERMITTENTEXPLOSIVE COMBUSTION AND IMPULSE TURBINE [76] Inventor: Harlow B. Grow,16530 Chattanooga Place, Pacific Palisades, Calif. 90272 [22] Filed:Oct. 30, 1973 [21] Appl. No.: 411,092

[52] US. Cl. 60/247; 60/261; 60/3975;

60/3981 [51] Int. Cl. F02k 3/10 [58] Field of Search 60/247, 248, 249,261,

60/39.7639.81, 39.65, 39.75, 39.74 R; 102/24 HC; 431/1 2,748,564 6/1956Marchal et a1 60/248 2,950,593 8/1960 Rae 60/261 X 3,138,920 6/1964Reichert 60/3975 3,618,319 11/1971 Kydd 60/3974 R OTHER PUBLICATIONSClark et al., Behavior of Metal Cavity Liners in Shaped ExplosiveCharges, Amer. Inst. of Mining &

[ Aug. 12, 1975 Metallurgical Engineers, 1947, Pub. No. 2158, pp, 1-3.

Primary Examiner-William L. Freeh Assistant Examiner--Robert E. Garrett[57] ABSTRACT A thrust producing gas turbine adapted for aircraftpropulsion at both moderate and especially high thrust values, andcharacterized by intermittent explosive burning of fuel into the workingfluid and the directing thereof into an impulse turbine for the partialabsorption of energy by means of a decrease in velocity and for theconservation of pressure for continued and subsequent effective burning.The engine embodiment comprises a compressor section followed by anozzlebumer section shaped for the explosive high pressure high velocitydischarge of working fluid into the turbine section for the absorptionof operating energy and conservation of pressure, and through which theworking fluid exhausts through stator blades that reestablish axial flowand thrust; the working fluid completing its combustion and/or beinginjected with fuel for efficient after-burning. The compressor air isstratified as it passes through the nozzle-burner section for coolingeffect and to supply the products of combustion for supporting completeprimary and secondary burning.

18 Claims, 3 Drawing Figures IMPULSE ROTOR \NFULSE STA-r0:

1 REACTION MOTOR EMPLOYING INTERMITTENT EXPLOSIVE COMBUSTION AND IMPULSETURBINE BACKGROUND The practice of jet propulsion involves the actionequals reaction principle and momentum or the quantity of motion definedas the product of the mass of a body and its velocity or M mv; m beingmass and v being velocity. The gas-turbine engine as it is employed forpurposes of jet propulsion involves these principles wherein the airintake condition is mvl and the exhaust condition is mv2 all as a resultof energy being applied by combustible fuels and operation of fluiddynamics dependent upon the energy transfer between a working fluid anda rotor.

Internal combustion jet engines are characterized by convergent anddivergent fluid flow passages embodied sequentially in a compressorsection, a combustion section, and a turbine section; it being a generalobject to provide thrust as efficiently as possible, or a difference inmass velocity between the intake and exhaust of the engine. Althoughmany variations of turbine design have been employed in aircraftpropulsion, the velocity turbine is employed exclusively to absorboperating energy from the working fluid, characterized by nozzles andexpansion blading and consequent increased velocity with a correspondingdrop in pressure as the working fluid moves rearward therethrough.Conventional jet turbines are therefore comprised of stationary bladesor nozzles through which the working fluid expands before flowingthrough the rotor blades which react and which may also be an expansivenozzle or nozzles; stage by stage as circumstances require. Such avelocity-reaction turbine is to be distinguished from the pureimpulse-reaction turbine with which this invention is particularlyconcerned and wherein the working fluid arrives at the nozzle entrancewith the approach velocity C V called the carry-Over velocity inmultistage turbines in the first stage of which C V is so small that itmay be neglected and the fluid may be assumed to start from rest.Significantly however, velocity is reduced commensurate with energyabsorption while P (entry pressure P at the nozzle discharge into thebucket or moving blade is conserved through the reaction function ofturning or redirecting the working fluid, and to the end that P P(outlet pressure P at the discharge from said buckets or moving blades.And it is the turning and/or redirection and the phenomenon that P Pwhich is advantageously employed in practicing may invention ashereinafter disclosed.

Fluid dynamics, especially when involved with explosive forces, involveswhat has become known as the Munroe effect by which expanding gases arefocused, so to speak, for efficient control over the work to beperformed thereby. In practice, explosive charges are directed by meansof this principle, either by shaping the charge or the chamber fromwhich it is to eminate. Similarly, it is an object of this invention toshape the nozzle section divergently for a most forceful impingement orhigh velocity contact of the working fluid upon the turbine blades;bearing in mind that it is the advantageous effect of momentum which isof primary concern, and that high pressure P and high velocity V are thecontrolling factors of efficient turbine operation.

Heat control is an important factor in the operation of internalcombustion jet engines, there being various ways to limit thetemperature of the working fluid as it passes through the turbineblading. For example, only part of the compressed air ispassed throughthe burner cans of the conventional jet engine, the remainder passingtherearound to restrict the temperature of the working fluid to a valuethat the turbine blading will withstand; and part of the air flow isbypassed, as for instance fan-jet engines. In this regard, it is anobject of this invention to separate the flow of working fluid withinthe nozzle section of a jet engine or gas-turbine. With this invention,the nozzle section is comprised of rearwardly divergent annular burnersarranged concentrically one within the other with an annular coolantduct therebetween.

Maximum thrust from jet engines is most often derived fromafter-burning, which is wasteful with conventional engines utilizingvelocity-reaction turbines having lowered exhaust pressures. It has beenlong recognized that high pressures are required for efficient burningof fuel, and conventional turbine engines are sufficiently efficientonly within the turbine section and are necessarily deficient, for thereasons advanced above, in the tail-pipe or exhaust section whereafterburning occurs. Therefore, it is an object of this invention toprovide a jet propulsion engine of the character thus far referred tothat maintains high pressure in the tail-pipe or exhaust section, forefficient after-burning. To this end, therefore, 1 have employed theimpulse turbine wherein P P the inlet and outlet pressures related ashereinabove set forth. Consequently, the tailpipe pressure after theturbine is at an efficient value nearly as great as the most efficientpressure preceding entry through the turbine blading.

Present day jet propulsion engines rely entirely upon the differencebetween M1 and M2 for thrust, as derived from increased velocity of theworking fluid; simply an increase in velocity. In contradistinction, itis an object of this invention to internally accelerate working fluid ofreduced velocity thereby effecting the difference in momentum M. ln thisregard, the characteristics of the impulse turbine are used to advantageto change axial flow into a vector or helical flow and converting aportion of the working fluid energy into rotative power, and duringwhich process the aircraft sees a substantial reduction in and/or to anextent no rearward thrust, and following which stator blading redirectsthe Working fluid axially to restore axial thrust at P P It is to beunderstood that the initial directional change from axial tocircumferential flow is less than percent as the working fluid passesfrom the burner-nozzles and into the turbine blading. By reaction of theworking fluid with impulse turbine blading, there is a redirecting ofthe working fluid with or without subsequent stator blading, and to theextent that the flow is directed circumferentially with a temporarynegation of forward thrust. However, by turbine blade design and/or byemploying subsequent stator blading, the remaining working fluid energyis reestablished as axial thrust and conserved as a propulsive thrust.Thus propulsive thrust is maintained at a reduced velocity and highpressure commensurate with the energy remaining in the working fluidafter removing a portion of the energy as power.

Conventional jet propulsion turbine engines have continuous burningsubject to flame-out and heat dissipation problems, it being an objectof this invention to provide highly efficient intermittently explosivehigh pressure burning ahead of the turbine and with coolant fluidadmixed therewith so as to control nozzle temperature of the workingfluid. With the present invention, intermittent fuel injection isprovided within the confines of the divergent nozzle-burner sectionshaped so as to gain the Monroe effect, and combined with ignition meanstherefor; whereby high velocity accompanies said high pressure forthrusting the working fluid through the turbine blading. Still further,the impulse blading that I employ conserves the pressure as P P forefficiently effective combustion aft of the turbine.

It is a general object of this invention to combine the aforementionedobjectives into a coordinated concept referred to herein as an aerothrust engine or the like, and which follows accepted engineeringprinciples for the most efficient operation thereof at both moderate andhigh thrust values.

DRAWINGS The various objects and features of this invention will befully understood from the following detailed description of the typicalpreferred form and application thereof, throughout which descriptionreference is made to the accompanying drawings, in which:

FIG. 1 is a longitudinal sectional view illustrating an aero thrustengine embodying the features of the present invention.

FIG. 2 is an enlarged detailed sectional view of the nozzle-burnersection of the engine shown in FIG. 1, and

FIG. 3 is an enlarged detailed section taken as indicated by line 3-3 ofFIG. 2.

PREFERRED EMBODIMENT The gas turbine engine of the present invention cantake various forms in each or all of its three major sections; namelythe compressor section X, the nozzleburner section Y and the turbinesection Z. Generally, the engine is of axial flow design wherein theintake condition is mvl and wherein the exhaust condition is mv2 Thecompressor section involves means to pressurize the flow of workingfluid therethrough and is characterized by a powered rotor, either ofthe centrifugal or axial flow type or a combination thereof; statorblading being provided as circumstances require in order to gain thedesired axial or helical flow as may be necessary. The nozzle-burnersection Y can vary according to the concept herein disclosed,characterized by its shape for the support of intermittent explosivecombustion induced therein. In practice, the nozzles are established bystator blading disposed so as to direct the propulsive thrust axiallywith accelerating velocity as combustion occurs therein. The turbinesection Z can also vary according to the concept herein disclosed,characterized by impulse blading or that type of turbine blading inwhich P P there being a reduction in velocity as energy is absorbed tocreate operating power used to drive the rotor of the compressor sectionX and the various accessories as may be required.

Referring now to the drawings, I have shown an axial flow compressorsection X that involves a plurality of stages of rotor and stator bladesand 11 that receive the intake air and reduce the volume thereof toincrease the pressure as it moves as working fluid therethrough. Theflow passage 12 is an annulus of large capacity at the intake and ofreduced capacity at the discharge into the nozzle-burner section Y. Arotor 13 carries the blades 10 while a stator 14 in the form of a caseand frame carries the blades 1 1, there being bearings 15 that journalthe rotor within the stator. It is to be understood that the compressorsection X pressures the working fluid to the highest pressurepermissible for the efficient primary and secondary burning to besubsequently accomplished, as will be described.

The nozzle-burner section Y forms an axial flow continuation of thecompressor section, and in accordance with this invention utilizes theMunroe effect, so to speak, for accelerating the working fluid in orderto increase velocity. The axial flow point at which P is established mayvary with engineering design, as it may be early at the juncture ofdischarge from the compressor section X into the nozzle-burner sectionY, or it may be later at any point up to the release of working fluidinto the turbine blades of turbine section Z. In any case, there is apoint in the axial flow where P, is established preceding entry of theworking fluid into and through the turbine blades.

In carrying out this invention, the burning of fuel in the working fluidis stratified for separation thereof from coolant fluid advantageouslyused for heat control and/or subsequent burning. Thus, there are atleast two layers of working fluid, one concentrically disposed withinthe other in annular strata established by a diffuser. In the engineillustrated there is an outer burner strata 20 and an inner burnerstrata 21, in which case there is at least one diffuser ring. Accordingto this in vention, a portion of the working fluid is separated out ascoolant and accordingly there is a pair of concentric diffuser rings 22and 23 between which there is an intermediate coolant strata 24 whichfolds rearwardly with increased velocity during the explosivecombustion. In the preferred form, therefore, the strata 20 and 21 areconcentric cylinders of divergent conical form having interfaceengagement with strata 24 rearward of the diffuser rings 22 and 23, thestrata 24 being of right cylinder form with the rings and of divergentconeshape rearward thereof.

The diffuser rings 22 and 23 are shaped so as to focus the burning offuel rearwardly, and they are characteristically divergent so as toinduce the rearward expansion and acceleration of the working fluid.Accordingly, the strata 20 and 21 are of expanding cross section as theyextend rearward, thereby to induce the expansive burning of fueltherein, the rings 22 and 23 terminating in sharp trailing edgessubstantially forward of the nozzle discharge immediately ahead of theturbine blading (see FIG. 2). Structural support for the diffuser rings22 and 23 is provided by radial struts 19 also in the form of statorvanes advantageously formed with fluid directive walls and terminatingin sharp trailing edges (see FIG. 2). In practice, a substantial portionof the nozzle-burner section Y is devoted to a burner chamber 25 forefficient directive explosion and for the comingling of gases whereby acooling effect takes place or is initiated ahead of the turbine blading.By properly proportioning the amount of coolant, a continued coolingeffect is produced within and for the protection of the turbine blading.

In accordance with this invention, regulatory fuel injection means andregulatory fuel ignition means are provided for introducing energy intothe working fluid. Each strata of working fluid designated for burning,such as layers and 21, are provided with fuel injectors 26 indicated asnozzles, and from which fuel is intermittently supplied from a fuel pumpmeans 27 and sprayed into the separated strata 20 and 21 respectively.The injectors 26 open into the strata 20 and 21 at either or both theinner and outer walls of the structure forming the same, and they arecircumferentially disposed at a frequency comensurate with the densityof fuel injection desired. In practice, the fuel injection means islocated at the forward end of the annular passages formed by thediffuser rings 22 and 23.

The regulatory fuel ignition means is shown as one or more ignitors 30,such as spark or glow plugs, located rearward of the fuel injectors.That is, ignition follows injection, there being at least one ring ofignitors 30 rearward of each ring of injectors 26, and preferably aseries of axially spaced rings of ignitors 30, 30', etc, adapted toprogressively ignite the injected fuel. The rings of ignitors 30 areaxially spaced as indicated and one ring thereof disposed so as toassume ignition within chamber 25.

In carrying out this invention, there is regulatory timed sequence (atintervals) of combustion by synchronizing the operation of means F and lin order to control power output by means of the frequency ofexplosions, each of which occurs at efficient pressure and velocity. Thefuel is intermittently injected at timed intervals by distribution fuelpump means 27, there being any number of patterns of injection sequencedependent upon the number and arrangement of said injectors 26. Andaccordingly, the injected fuel charges are subsequently ignited byelectrical discharge distributor means 28 of the ignition means l, timedso as to follow the injection with a spark or the like that isdischarged for the intermittent igniton of the fuel charges, as they areinjected.

The compressor section X and nozzle-burner section Y cooperate toprovide intermittent explosive burning of fuel into the working fluid soas to increase the energy and the discharge of said working fluid intothe turbine section Z at a high velocity and an efficient pressure. Thevelocity and resultant thrust can be varied as required, by controllingthe amount of fuel injected, as by throttling or by frequency andduration of injection. Thus, the working fluid that moves intermediatethe fuel charges and explosive burning thereof has a heat controleffect; the intermediate strata of coolant 24 has its continuing heatcontrol effect; and all to the end that working fluid temperature at thedischarge of the nozzle-burner is under determinative control, and whichis restricted to be within the temperature limit that the turbineblading will withstand.

The turbine section Z includes the turbine wheel and tail pipe withinwhich continued primary or secondary burning is efficientlyaccomplished. As shown, the blading of turbine wheel 35 forms an annularcontinuation of the nozzle chamber 25 for the axial flow of workingfluid therethrough. ln accordance with this invention, the blades 36 areof the impulse typeby which velocity is decreased in order to absorbenergy, while pressure is maintained. Thus, mvl is substantially greaterthan mv2 while P substantially equals P Consequently, the efficientburning pressure at the nozzle-burner discharge is maintainedimmediately rearward of the turbine wheel 35 for the advantageoussecondary burning of fuel. In practice, complete primary burning withinthe axial confines of chamber 25 may be preferred, however, with theapplication of coolant 24 continued primary burning and/or thecompletion thereof, from chamber 25, can occur as the working fluidpasses through the turbine wheel 35, and said con tinued burning issupported by the pressure value P P As shown, secondary burner injectors37 are provided immediately rearward of the turbine wheel 35 and in theefficient pressure zone P P of the tail pipe 40. The divergent annulusof the tail pipe surrounding the tail cone 41 provides for the furtherexpansion of secondary burning fuel, and with proper design, anefficient combustion pressure is maintained and a continued drop inpressure permitted thereafter for exhaust of working fluid into theatmosphere.

In accordance with this invention, the change in direction of workingfluid as caused by reaction thereof through the impulse turbine bladingsubtracts energy from the axial flow in proportion to the power providedfor the compressor and auxillary operation as compared with the totalenergy of the working fluid delivered by the nozzle-burner. To thisproportionate extent, therefore, the aircraft sees a reduced thrust as aresult of the lateral (helical) deflection of the working fluid leavingthe turbine blades 36. The velocity at P is also commensurately reduced.However, this thrust is but temporarily reduced since it is regained bystator blades 38 immediately rearward of turbine blades 36 and thatredirect the working fluid to flow axially and to the end that the totalenergy (momentum) remaining in the working fluid is seen by the aircraftas propulsive thrust. Therefore, it will be seen that there is theconservation of both momentum and pressure for the support of efficientcontinued burning and/or secondary burning for the increased thrustafforded thereby and which is determinative of the ultimate thrust to bedeveloped by the engine.

Having described only a typical preferred form and application of myinvention, 1 do not wish to be limited or restricted to the specificdetails herein set forth, but wish to reserve to myself anymodifications or variations that may appear to those skilled in the art.

Having described my invention, I claim:

l. A moderate thrust producing gas-turbine including; a compressorsection having a working fluid inlet and means discharging compressedworking fluid axially at a pressure adapted to support combustion, anozzle-burner section receiving the axial discharge from said compressormeans and comprised of a rearwardly divergent passage conducting theworking fluid and with its nozzle area greater than its reception areafrom the compressor section and having intermittently explosivecombustion means for expansion and axial velocity increase of saidworking fluid, and a turbine section receiving the working fluid fromthe nozzleburner section at said axial velocity increase and having aturbine with energy absorbing velocity decreasing and pressureconserving blading driving said compressor and discharging remainingworking fluid axially rearward as propulsive thrust.

2. A thrust producing gas-turbine as set forth in claim 1, wherein thecombustion means comprises injection and intermittent ignition meanscharging fuel into the nozzle-burner for theexplosive and rearwardlyfocused combustion thereof.

5. A thrust producing gas-turbine as set forth in claim' 1, wherein thecombustion means comprises intermittent injection means charging thefuel and an axial spaced ignitor means for the explosive combustionthereof.

6. A thrust producing gas-turbine as set forth in claim 1, wherein thecombustion means comprises intermittent injection means charging thefuel and a series'of axially spaced ignitor means for the progressiveexplosive combustion thereof.

7. A thrust producing gas-turbine as set forth in claim 1, wherein thenozzle-burner section has concentrically spaced diffuser rings disposedtherein stratifying the working fluid into spaced layers for separatingcombustion with a coolant layer therebetween, and wherein intermittentinjection means introduces fuel into said spaced layers and axiallyspaced ignitors combust the explosive charges for rearward accelerationand comingling with said coolant layer.

8. A high thrust producing gas-turbine including; a compressor sectionhaving a working fluid inlet and means discharging compressed workingfluid axially at a pressure adapted to support primary combustion, anozzle-burner section receiving the axial discharge from said compressormeans and comprised of a rearwardly divergent passage conducting'theworking fluid and with its nozzle area greater than its reception areafrom the compressor section and having a primary combustion means forexpansion and velocity increase of said working fluid, a turbine sectionreceiving the working fluid from the nozzle-burner section at said axialvelocity increase and having a turbine with energy absorbing velocitydecreasing and pressure conserving rotor and stator blading driving saidcompressor and for maintaining working fluid pressure adapted to supportsecondary combustion axially rearward thereof, and secondary combustionmeans for the further expansion and velocity increase of said workingfluid discharging as propulsive thrust.

11. The high thrust producing gas-turbine as set forth in claim 8,whereinv the turbine section rotor blading is of impulse configurationfor the conservation of pressure therethrough. i

12. The high thrust producing gas-turbine as set forth in claim 8,wherein the turbine se'ction stator blading is of impulse vconfigurationfor the conservation of pressure therethrough.

13. The high thrust producing gas-turbine as set forth in claim 8,wherein the turbine s ection rotor blading and stator blading rearwardthereof are of impulse configuration for the conservation of pressuretherethrough.

14. The high thrust producing gas-turbine as set forth.

in claim 8, wherein the turbine'section stator blading is of impulseconfiguration for the conservation of pressure therethrough, and whereinthe secondary combus- 9. The high thrust producing gas-turbine as setforth in claim 8, wherein the secondary combustion means comprises fuelinjection means charging fuel into the working fluid rearward of saidturbine blading.

10. The high thrust producing gas-turbine as set forth in claim 8,wherein the secondary combustion means comprises fuel injection andignition means charging fuel into the working fluid rearward of saidturbine blading.

tion means comprises fuel injection and ignition means charging fuelinto the pressure conserved area of the working fluid. j

15. The high thrust producing gas-turbine as set forth in claim 8,wherein the turbine section rotor blading and stator blading rearwardthereof are of impulse configuration for the conservation of pressuretherethrough, and wherein the secondary combustion means comprises fuelinjector and ignition means charging fuel into the pressure conservedarea of theworking 16. A high thrust producing gas-turbine including; acompressor section'ha'ving a working fluid inlet and meansdischargingcompressed working fluid axially at a pressure adapted tosupport primary combustion, a nozzle-burner section receiving the axialdischarge from said compressor means and. comprised of a rearwardlydivergent passageconducting the working fluid and with its nozzle areagreater than its reception area from the compressor section and havingintermittently explosive primary combustion means for expansion andaxial velocity increase of said working fluid, a turbine sectionreceiving the working fluid from the nozzle burner section at said axialvelocity increase and having a turbine with energy absorbing velocitydecreasing and pressure conserving blading driving said .compressor andformaintaining working fluid pressure adapted to support secondarycombustion axially rearward thereof, and secondary combustionmeans'forthe further expansion and velocity increase of said workingfluid discharging as propulsive thrust.

17. The high thrust producing gas-turbine as set forth in claim 16,wherein the nozzle-burner section includes a diffuser ring stratifyingthe working fluid into separate layers for combustion and coolingthereof respectively. 7

18. The high thrust producing gas-turbine as set forth in claim 16,wherein the nozzle-burner section includes spaced diffuser ringsstratifying the working fluid into spaced layers for separate combustionwith a coolant layer therebetween.

1. A moderate thrust producing gas-turbine including; a compressorsection having a working fluid inlet and means discharging compressedworking fluid axially at a pressure adapted to support combustion, anozzle-burner section receiving the axial discharge from said compressormeans and comprised of a rearwardly divergent passage conducting theworking fluid and with its nozzle area greater than its reception areafrom the compressor section and having intermittently explosivecombustion means for expansion and axial velocity increase of saidworking fluid, and a turbine section receiving the working fluid fromthe nozzle-burner section at said axial velocity increase and having aturbine with energy absorbing velocity decreasing and pressureconserving blading driving said compressor and discharging remainingworking fluid axially rearward as propulsive thrust.
 2. A thrustproducing gas-turbine as set forth in claim 1, wherein the combustionmeans comprises injection and intermittent ignition means charging fuelinto the nozzle-burner for the explosive and rearwardly focusedcombustion thereof.
 3. A thrust producing gas-turbine as set forth inclaim 1, wherein the nozzle-burner section includes a diffuser ringwithin said rearwardly divergent passage and stratifying the workingfluid into separate layers for combustion and cooling thereofrespectively.
 4. A thrust producing gas-turbine as set forth in claim 1,wherein the nozzle-burner section includes spaced diffuser rings withinsaid rearwardly divergent passage and stratifying the working fluid intospaced layers for separate combustion with a coolant layer therebetween.5. A thrust producing gas-turbine as set forth in claim 1, wherein thecombustion means comprises intermittent injection means charging thefuel and an axial spaced ignitor means for the explosive combustionthereof.
 6. A thrust producing gas-turbine as set forth in claim 1,wherein the combustion means comprises intermittent injection meanscharging the fuel and a serIes of axially spaced ignitor means for theprogressive explosive combustion thereof.
 7. A thrust producinggas-turbine as set forth in claim 1, wherein the nozzle-burner sectionhas concentrically spaced diffuser rings disposed therein stratifyingthe working fluid into spaced layers for separating combustion with acoolant layer therebetween, and wherein intermittent injection meansintroduces fuel into said spaced layers and axially spaced ignitorscombust the explosive charges for rearward acceleration and cominglingwith said coolant layer.
 8. A high thrust producing gas-turbineincluding; a compressor section having a working fluid inlet and meansdischarging compressed working fluid axially at a pressure adapted tosupport primary combustion, a nozzle-burner section receiving the axialdischarge from said compressor means and comprised of a rearwardlydivergent passage conducting the working fluid and with its nozzle areagreater than its reception area from the compressor section and having aprimary combustion means for expansion and velocity increase of saidworking fluid, a turbine section receiving the working fluid from thenozzle-burner section at said axial velocity increase and having aturbine with energy absorbing velocity decreasing and pressureconserving rotor and stator blading driving said compressor and formaintaining working fluid pressure adapted to support secondarycombustion axially rearward thereof, and secondary combustion means forthe further expansion and velocity increase of said working fluiddischarging as propulsive thrust.
 9. The high thrust producinggas-turbine as set forth in claim 8, wherein the secondary combustionmeans comprises fuel injection means charging fuel into the workingfluid rearward of said turbine blading.
 10. The high thrust producinggas-turbine as set forth in claim 8, wherein the secondary combustionmeans comprises fuel injection and ignition means charging fuel into theworking fluid rearward of said turbine blading.
 11. The high thrustproducing gas-turbine as set forth in claim 8, wherein the turbinesection rotor blading is of impulse configuration for the conservationof pressure therethrough.
 12. The high thrust producing gas-turbine asset forth in claim 8, wherein the turbine section stator blading is ofimpulse configuration for the conservation of pressure therethrough. 13.The high thrust producing gas-turbine as set forth in claim 8, whereinthe turbine section rotor blading and stator blading rearward thereofare of impulse configuration for the conservation of pressuretherethrough.
 14. The high thrust producing gas-turbine as set forth inclaim 8, wherein the turbine section stator blading is of impulseconfiguration for the conservation of pressure therethrough, and whereinthe secondary combustion means comprises fuel injection and ignitionmeans charging fuel into the pressure conserved area of the workingfluid.
 15. The high thrust producing gas-turbine as set forth in claim8, wherein the turbine section rotor blading and stator blading rearwardthereof are of impulse configuration for the conservation of pressuretherethrough, and wherein the secondary combustion means comprises fuelinjector and ignition means charging fuel into the pressure conservedarea of the working fluid.
 16. A high thrust producing gas-turbineincluding; a compressor section having a working fluid inlet and meansdischarging compressed working fluid axially at a pressure adapted tosupport primary combustion, a nozzle-burner section receiving the axialdischarge from said compressor means and comprised of a rearwardlydivergent passage conducting the working fluid and with its nozzle areagreater than its reception area from the compressor section and havingintermittently explosive primary combustion means for expansion andaxial velocity increase of said working fluid, a turbine sectionreceiving the working fluid from the nozzle-burner section at said axialvelocIty increase and having a turbine with energy absorbing velocitydecreasing and pressure conserving blading driving said compressor andfor maintaining working fluid pressure adapted to support secondarycombustion axially rearward thereof, and secondary combustion means forthe further expansion and velocity increase of said working fluiddischarging as propulsive thrust.
 17. The high thrust producinggas-turbine as set forth in claim 16, wherein the nozzle-burner sectionincludes a diffuser ring stratifying the working fluid into separatelayers for combustion and cooling thereof respectively.
 18. The highthrust producing gas-turbine as set forth in claim 16, wherein thenozzle-burner section includes spaced diffuser rings stratifying theworking fluid into spaced layers for separate combustion with a coolantlayer therebetween.